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AE 443S - TEAM 4
APRIL 16TH, 2015
FDR PRESENTATION
2
THE TEAM
Jay Mulakala
Lead Systems
Engineer
Samip Shah
ADCS Systems
Engineer
Bentic Sebastian
Power and Thermal
Systems ...
PLAN OF ACTION
3
Mission
Summary
Vehicle
Systems
Overview
Mission
Architecture
Risk Analysis
Mission
Costs
MISSION SUMMARY
JAY MULAKALA
THE CRITERIA
5
 The OTV will be stationed in 400
km AMSL circular LEO with 28°
inclination.
 The OTV payload capability ...
CENTURION
6
CENTURION
7
THE CENTURION
8
TIMELINE
9
TIMELINE
10
VEHICLE SYSTEMS OVERVIEW
JAY MULAKALA
MISSION SYSTEMS
12
• Orbital Systems
• Spacecraft Propulsion Systems
• Structural Definition
• Communications
• ADCS
• Spa...
DEREK AWTRY
ORBITAL SYSTEMS
DESIGN PROCESS
 Orbital Requirements
 Maximum transfer time of 6 days to L1 and L2
 Orbit around L1/L2 for at least 30 ...
CONCEPT DEVELOPMENT
15
 Trajectory to L1
 Different halo orbits were simulated by varying the z-amplitude on the
Earth-M...
CONCEPT DEVELOPMENT
16
Trajectory to L1 using STK
CONCEPT DEVELOPMENT
17
 Trajectory to L2
 Different halo orbits were simulated by varying the z-amplitude on the
Earth-M...
CONCEPT DEVELOPMENT
18
Trajectory to L2 using STK
CONCEPT DEVELOPMENT
19
 Station-Keeping Analysis
 A station-keeping maneuver was added for each orbit simulated
 ΔV1 is...
CONCEPT DEVELOPMENT
20
 Orbital Maintenance In LEO
 Refueling in LEO will take place at the end of the 4-5 month period ...
CONCEPT DEVELOPMENT
21
 Equation derived by NASA for ΔV savings
 Initially used for the Martian atmosphere, can be
expan...
BENJAMIN WILSON
PROPULSION SYSTEMS
DESIGN PROCESS
Main Propulsion System
Requirements • For a round trip to L2 provide a total ∆V of 8.31 km/s
• Minimize fue...
CONCEPT DEVELOPMENT: MAIN PROPULSION
Type Thruster Thrust
[N]
Specific Impulse [S] Fuel to L2
[kg]
Ion
Aerojet NEXT 0.235 ...
Escort BNTR System Characteristics
System
Number of Units 3
Total System Mass 6675 kg
Thrusters
Thrust Per Unit 111,200 N
...
26
NUCLEAR REACTOR SAFETY
 Shielding to reduce exposure to less than 1 REM/year
 Reactor shutdown while docking and refueli...
MAIN PROPULSION SYSTEM PROPELLANT AND TANKAGE
 Liquid Hydrogen propellant
 Ensures Lowest Fuel mass
 700m3 composite fu...
PROPELLANT MASS CALCULATION TO L2 [ISP=911S]
Stage Description ∆V [m/s] Time Elapsed
Mi (Includes
Payload) [kg]
Mpi Propel...
CONCEPT DEVELOPMENT: ATTITUDE CONTROL PROPULSION
Type Engine Fuel Isp [s] Thrust [N] Propellant
Mass[ kg]
Ion Aerojet NEXT...
ACS PROPELLANT AND TANKAGE
Tank Propellant Volume
(L)
Mass
(kg)
Tanks
Required
MOOG GEO Sat. Hydrazine 220 27 4
ATK 80505-...
YU GUAN
STRUCTURAL DEFINITION
CONCEPT DEVELOPMENT
33
• Systems module
• ADCS sensors
• Power and Thermal units
• Propellant tank
• Thermal shielding for...
SYSTEMS MODULE
34
CRYOGENIC PROPELLANT TANK
35
• Internal volume of 625 cubic meters
• Aluminum lithium thermal shielding
PROPULSION SYSTEM
36
STRUCTURAL COMPONENTS
37
CYCOM 5320-1 epoxy resin
systemfrom Cytek Inc. [12]
PAMG-XR1 aerospace grade
aluminum honeycomb f...
RADIATOR DESIGN
38
Hinged radiator used in systems
module. [14]
Deployable radiator installed in
propulsion system. [15]
MASS ESTIMATION
39
STRUCTURAL TESTING
40
YU GUAN
COMMUNICATIONS
DESIGN PROCESS
 Network Selection
 Network ideally suited for needs of the Centurion
 Band Selection
 Allow for uninte...
CONCEPT DEVELOPMENT – NETWORK SELECTION
 NEN(Near Earth Network)
performance is comparable to
DSN(Deep Space Network) at ...
CONCEPT DEVELOPMENT – FREQUENCY BAND SELECTION
Several bands
 S-band
 Low frequency
 Low data rates
 Lower pointing re...
CONCEPT DEVELOPMENT – RADIOMETRIC TRACKING
Doppler
 Provides velocity estimates
 Orbital maneuvers
 ΔV calculations
Ran...
CONCEPT DEVELOPMENT
Parabolic reflector
X & S band
Low gain as backup
Foldable design
46
CRITICAL DESIGN ISSUES
 Structural Critical Issues
- Thermal Cycling
- Material Failure
 Communication Critical Issues
-...
SAMIP SHAH
ATTITUDE DETERMINATION & CONTROL SYSTEMS
DESIGN PROCESS
 Sensor Selection
 High attitude sensing requirements for docking and refueling
 Actuator Selection
 Hi...
CONCEPT DEVELOPMENT – SENSOR SELECTION
Star Trackers
 Absolute attitude sensor
 Highest accuracy
 Low update rate
Selec...
CONCEPT DEVELOPMENT – SENSOR SELECTION (CONT.)
Inertial Measurement Units
 Relative attitude sensor
 Prone to drift
 Hi...
CONCEPT DEVELOPMENT – SENSOR SELECTION (CONT.)
Sun Sensors
 Useful for solar array pointing
 Relatively inexpensive
Sele...
CONCEPT DEVELOPMENT – SENSOR SELECTION (CONT.)
Proximity Sensor
 Vital for autonomous docking maneuvers
 Enables accurat...
CONCEPT DEVELOPMENT – ACTUATOR SELECTION
Attitude Control
Thrusters
 Provide both large and
fine attitude
adjustments
 1...
CONCEPT DEVELOPMENT – ACTUATOR SELECTION (CONT.)
Control Moment Gyroscopes
 Common on large spacecraft
 Easily mitigates...
56
CONCEPT DEVELOPMENT – CONTROL METHODS
Onboard Processing
 Comprehensive processing is required to control attitude and po...
CONCEPT DEVELOPMENT – CONTROL METHODS (CONT.)
58
SOURCE LINES OF CODE
Computer Software Component SLOC
Executive 1,000
Communications 2,000
Attitude/Orbit Sensor Processin...
BENTIC SEBASTIAN
SPACECRAFT POWER MANAGEMENT
DESIGN PROCESS
 Four performance requirements
 Supply power for all instruments onboard.
 Provide suitable radiation sh...
CONCEPT DEVELOPMENT- POWER GENERATION AND DISTRIBUTION
62
 Main source of power is BNTRS
 Three Nuclear Reactors, produc...
CONCEPT DEVELOPMENT- POWER GENERATION AND DISTRIBUTION
 Peak voltage = 50V for Surrey/Rigel-L star tracker
 Peak power i...
POWER GENERATION AND DISTRIBUTION
64
CONCEPT DEVELOPMENT – POWER STORAGE
 Requirements for batteries
 Light
 Rechargeable
 High energy density
 Long disch...
CONCEPT DEVELOPMENT – POWER STORAGE
66
Technology Specific
Density(Wh/kg)
Energy
Density(Wh/l)
Operating temp.
Range(C)
De...
CONCEPT DEVELOPMENT – POWER STORAGE
67
 Design choice:
 Eight Quallion QL075KA batteries, 72Ah, 3.6V
 Eight additional ...
CONCEPT DEVELOPMENT – RADIATION SHIELDING
 Material for Radiation
Shielding:
 Aluminum
 Material Thickness
 3cms
 Shi...
CONCEPT DEVELOPMENT – EMERGENCY MODE
 Two solar panels will generate additional energy during emergency conditions
 Mini...
BENTIC SEBASTIAN
SPACECRAFT THERMAL SYSTEMS
DESIGN PROCESS
 Control temperature within
 Nuclear Reactors
 Fuel Tank
 Systems Module
 Solar Panels and Radiators
7...
CONCEPT DEVELOPMENT – THERMAL CONTROL OF NUCLEAR REACTORS
72
Brayton Cycle of ESCORT System [29]
 Peak temperature of HeX...
CONCEPT DEVELOPMENT – THERMAL CONTROL OF NUCLEAR REACTORS
 Deployable radiators
designed by Lockheed
Martin.
 Total area...
CONCEPT DEVELOPMENT – THERMAL CONTROL OF FUEL TANK
74
Detailed wireframe of the OTV
 Cryogenic fuel tank made of Aluminum...
CONCEPT DEVELOPMENT – THERMAL CONTROL OF SYSTEMS MODULE AND
SOLAR PANELS
75
 Two hinged Radiators of size 2m by 0.5m will...
KEVIN LOHAN
LAUNCHING AND DOCKING
DESIGN PROCESS
 Launch vehicle
 OTV launch vehicle must be reliable
 Single launch
 Universal docking mechanism
 Carr...
CONCEPT DEVELOPMENT : LAUNCH VEHICLE
78
Launch Vehicle
Payload
Capacity (kg)
Launch Cost
(millions of $)
Falcon XX 140,000...
CONCEPT DEVELOPMENT : DOCKING SYSTEM
79
 International docking system standards (IDSS)
 Regulates where connections are ...
CONCEPT DEVELOPMENT : REFUELING
80
 Robotic Refueling
Mission (RRM)
 Modified Dextre Arm
 Successful test in
2013
Robot...
JAY MULAKALA
RISK ANALYSIS
TECHNOLOGY RISK ANALYSIS
82
TECHNOLOGY CONSEQUENCE POF SHORT CODE
Star Trackers 3 3 [1]
IMU's 3 3 [2]
Magnetometer 1 3 [3]...
OPERATIONAL RISK ANALYSIS
83
OPERATIONS CONSEQUENCE PO SHORT CODE
Political setbacks due to technologies 4 4 [1]
Use of nu...
84
RISK MITIGATION
Technology Risk Analysis
Operational Risk Analysis
Main Factors:
• Nuclear Thermal
Propulsion System
• ...
JAY MULAKALA
MISSION COSTS
COST ESTIMATION
86
FixedCosts
• Developmen
t
• Production
• Assembly
• $2.518
billion
ClientCosts
• Fuel
Transport
• Fuel ...
COST ESTIMATION
87
COST ESTIMATION
88
Centurion’s Liquid
Hydrogen Fuel
• 10 Missions
• 10 Falcon 9
Heavys
• $850 million
Conventional
Biprope...
COST ESTIMATION
89
Our Solution
• Delta IV
• Nuclear bi-
modal
propulsion
system
• Modified NASA
Docking System
Current So...
THE CENTURION
90
REFERENCES
REFERENCES
REFERENCES
COST ESTIMATION – FIXED COSTS
94
Centurion - OTV Presentation
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Centurion - OTV Presentation

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The Centurion Orbit Transfer Vehicle (OTV) was part of our Aerospace Engineering Senior Design project at the University of Illinois at Urbana-Champaign. It is equipped with the latest technologies, including a nuclear thermal propulsion system. The structure weighs 89,000 kg and is capable of transporting cargo to Lagrange points L1 or L2.

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Centurion - OTV Presentation

  1. 1. AE 443S - TEAM 4 APRIL 16TH, 2015 FDR PRESENTATION
  2. 2. 2 THE TEAM Jay Mulakala Lead Systems Engineer Samip Shah ADCS Systems Engineer Bentic Sebastian Power and Thermal Systems Engineer Yu Guan Structures Engineer Ben Wilson Propulsion Engineer Derek Awtry Orbital Systems Engineer Kevin Lohan Launch and Docking Engineer
  3. 3. PLAN OF ACTION 3 Mission Summary Vehicle Systems Overview Mission Architecture Risk Analysis Mission Costs
  4. 4. MISSION SUMMARY JAY MULAKALA
  5. 5. THE CRITERIA 5  The OTV will be stationed in 400 km AMSL circular LEO with 28° inclination.  The OTV payload capability shall be 50,000 lbs from LEO to EML1 and 15,000 lbs from EML1 to LEO.  The OTV must be capable to remain at EML1 or EML2 for at least 30 days.  Each transfer should not exceed 6 days.  The life of the OTV shall be 5 years and the OTV shall be capable of at least 10 missions to EML1 or EML2.
  6. 6. CENTURION 6
  7. 7. CENTURION 7
  8. 8. THE CENTURION 8
  9. 9. TIMELINE 9
  10. 10. TIMELINE 10
  11. 11. VEHICLE SYSTEMS OVERVIEW JAY MULAKALA
  12. 12. MISSION SYSTEMS 12 • Orbital Systems • Spacecraft Propulsion Systems • Structural Definition • Communications • ADCS • Spacecraft Power Management Systems • Spacecraft Thermal Systems • Launching and Docking Systems
  13. 13. DEREK AWTRY ORBITAL SYSTEMS
  14. 14. DESIGN PROCESS  Orbital Requirements  Maximum transfer time of 6 days to L1 and L2  Orbit around L1/L2 for at least 30 days  Initial Low Earth Orbit of 400 km and 28˚ inclination  Can consider Aerobraking  STK/Astrogator was used to determine trajectories to L1 and L2  Multiple trajectories were considered  The trajectories with the lowest ΔV was chosen for each Lagrange point  We will not consider aerobraking 14
  15. 15. CONCEPT DEVELOPMENT 15  Trajectory to L1  Different halo orbits were simulated by varying the z-amplitude on the Earth-Moon plane.  Each initial burn from LEO is 3.069 km/s, each burn back into LEO was 3.058 km/s, with a time of flight (TOF) there and back of 4.3-4.4 days Orbit Option Amplitude (km) 𝚫𝐕𝟏 (m/s) 𝚫𝐕𝟐 (m/s) Total 𝚫𝐕 (m/s) Halo Orbit (days) 1 5,000 620.017 644.108 7411.944 35.999 2 7,500 622.213 644.063 7421.159 36.146 3 10,000 625.070 647.151 7422.128 36.108 4 15,000 632.660 654.051 7446.725 36.079 5 20,000 642.570 663.845 7482.638 36.140 ΔV’s for different halo orbits
  16. 16. CONCEPT DEVELOPMENT 16 Trajectory to L1 using STK
  17. 17. CONCEPT DEVELOPMENT 17  Trajectory to L2  Different halo orbits were simulated by varying the z-amplitude on the Earth-Moon plane.  Each initial burn from LEO is 3.094 km/s, each burn back into LEO was 3. 099 km/s, with a TOF there of 5.3 days and a TOF back of 5.9 days Orbit Option Amplitude (km) 𝚫𝐕𝟏 (m/s) 𝚫𝐕𝟐 (m/s) Total 𝚫𝐕 (m/s) Halo Orbit (days) 6 5,000 1124.806 1018.230 8352.291 44.277 7 7,500 1122.052 1030.948 8372.584 45.089 8 10,000 1118.017 1003.735 8339.403 45.171 9 15,000 1106.847 988.601 8321.102 44.985 10 20,000 1097.024 974.496 8306.030 44.841 ΔV’s for different halo orbits
  18. 18. CONCEPT DEVELOPMENT 18 Trajectory to L2 using STK
  19. 19. CONCEPT DEVELOPMENT 19  Station-Keeping Analysis  A station-keeping maneuver was added for each orbit simulated  ΔV1 is the station-keeping burn after the first revolution  ΔV2 is the station-keeping burn after the second revolution Orbit Option 𝚫𝐕𝟏 (m/s) 𝚫𝐕𝟐 (m/s) Total 𝚫V (m/s) 1 9.586 9.952 19.538 2 13.588 13.069 26.657 3 7.352 14.37 21.722 4 7.834 24.048 31.882 5 10.311 37.805 48.116 Station-keeping maneuvers at L1 Station-keeping maneuver at L2Orbit Option 𝚫𝐕𝟏 (m/s) 𝚫𝐕𝟐 (m/s) Total 𝚫V (m/s) 6 8.386 6.922 15.308 7 8.729 16.92 25.657 8 9.412 14.139 23.551 9 12.765 18.727 31.492 10 17.744 22.860 40.604
  20. 20. CONCEPT DEVELOPMENT 20  Orbital Maintenance In LEO  Refueling in LEO will take place at the end of the 4-5 month period between missions  ΔV to get back to required orbit is 6.797 m/s  End of Life Summary  A payload will be taken on a one way mission to L1, and dropped off  Centurion will then maneuver to the stable EM-L4  ΔV from L1 to L4 is around 682 m/s  Final mission ΔV = 4.914 km/s
  21. 21. CONCEPT DEVELOPMENT 21  Equation derived by NASA for ΔV savings  Initially used for the Martian atmosphere, can be expanded for all celestial bodies [3]  ΔV = 𝐶 𝑑 𝑆 2𝑚 ρ0 2πμ 1+e e RT g Altitude (km) Atmospheric Density (kg/km^3) 𝚫V savings (m/s) 50 102700.0 1,026.4 60 30960.0 295.6152 80 18449.456 157.9763 100 560.276 4.7543 120 22.234 0.2563 140 3.839 0.552 ΔV savings for different periapsis altitudes
  22. 22. BENJAMIN WILSON PROPULSION SYSTEMS
  23. 23. DESIGN PROCESS Main Propulsion System Requirements • For a round trip to L2 provide a total ∆V of 8.31 km/s • Minimize fuel mass • Ensure safe operation for passengers • Reliable 23 Attitude Control Propulsion System Requirements • Provide the total ∆V required for attitude control over the entire operational lifetime of the Centurion • Provide a total ∆V of 10 m/s • Minimize fuel mass • Reliable
  24. 24. CONCEPT DEVELOPMENT: MAIN PROPULSION Type Thruster Thrust [N] Specific Impulse [S] Fuel to L2 [kg] Ion Aerojet NEXT 0.235 4100 8,700 Busek BHT-20k 0.807 2320 16,600 NASA NSTAR 0.094 3195 11,600 Bipropellant Aerojet CECE 111,000 465 183,00 Astrium Aestus 29,600 324 436,000 CALT YF-73 44,150 420 229,000 Monopropellant Aerojet MR-80B 3780 225 1,410,000 AMPAC MONARC 445 445 235 1,190,000 Nuclear Thermal NERVA XE 1,112,000 850 62,500 Escort 333,600 911 49,600 CIS NTR 111,600 941 47,500 24 [9] [10] [11] [12] [13] [14] [15] [16]
  25. 25. Escort BNTR System Characteristics System Number of Units 3 Total System Mass 6675 kg Thrusters Thrust Per Unit 111,200 N Specific Impulse 911 s Propellant Propellant Liquid Hydrogen Propellant Mass 49,959 kg Propellant volume 700 m3 Tank Material Composite Tank Boil Off 1% per month by mass Reactors Fission Material UO2-W Cermet (Uranium Dioxide in Tungsten matrix) Exhaust Temperature ~ 2700 Kelvin Power & Thermal Power Per Unit 25kW Power Cycle Brayton Cycle Power Fluid Helium Xenon Radiator Area 65m2 25
  26. 26. 26
  27. 27. NUCLEAR REACTOR SAFETY  Shielding to reduce exposure to less than 1 REM/year  Reactor shutdown while docking and refueling  No critical state before leaving the atmosphere  Not allowed to re-enter atmosphere until all fissile material has been used  Will be decommissioned at EML4 27
  28. 28. MAIN PROPULSION SYSTEM PROPELLANT AND TANKAGE  Liquid Hydrogen propellant  Ensures Lowest Fuel mass  700m3 composite fuel tank  Active and passive Thermal control to achieve ~1% boil off per month 28 NASA and Boeing’s 5.5 Meter cryogenic composite fuel tank [61] 𝐼𝑠𝑝 = 𝐴𝐶𝑓 𝑇𝑐 𝑀
  29. 29. PROPELLANT MASS CALCULATION TO L2 [ISP=911S] Stage Description ∆V [m/s] Time Elapsed Mi (Includes Payload) [kg] Mpi Propellant Spent[kg] 1 Departing LEO 3095 10 Minutes 87,745 25,682 2 Transit to L2 0 5.3 Days 62,063 42 3 Arrive at L2 1097 3 Minutes 62,021 7,165 4 Halo orbit 1 0 15 days 32,177 83 5 Halo Correction 1 18 6 Seconds 32,094 64 6 Halo Orbit 2 0 15 days 32,030 83 7 Halo Correction 2 23 6 Seconds 31,947 82 8 Halo Orbit 3 0 15 days 31,866 83 9 Depart L2 974 2 Minutes 38,588 3,986 10 Transit to LEO 0 6 days 34,602 25 11 Arrive at LEO 3099 4 Minutes 34,577 10,133 Final Mass = 24,444 𝑀 𝑝𝑖 = 47,428 29 𝑀 𝑝 = 𝑀𝑠𝑦𝑠(𝑒∆𝑉 𝐼𝑠𝑝 𝑔 𝑜 − 1)
  30. 30. CONCEPT DEVELOPMENT: ATTITUDE CONTROL PROPULSION Type Engine Fuel Isp [s] Thrust [N] Propellant Mass[ kg] Ion Aerojet NEXT Xenon 4100 0.235 13 Busek BHT-20k Xenon 2320 0.807 23 Cold Gas MOOG 58-118 Unknown 72 3.5 560 AMPAC SVT01 Xenon 45 0.05 900 Monopropellant AMPAC MONARC -90 Hydrazine 235 90 215 Aerojet MR-107N Hydrazine 232 109-296 220 Bipropellant EADS 10N NTO, MON-1, MON-3 and MMH 291 10 195 Aerojet R-1E MMH/NTO 280 111 190 30 [51] [52] [53]
  31. 31. ACS PROPELLANT AND TANKAGE Tank Propellant Volume (L) Mass (kg) Tanks Required MOOG GEO Sat. Hydrazine 220 27 4 ATK 80505-1 Any 134 16 4 Astrium OST 31/0 MON/MMH 235 16 4 31 Propellant Volume (L) Mass (kg) Mono methyl hydrazine (MMH) 83 73 Nitrogen tetra oxide (NTO) 81 117 ATK 80505-1[63]
  32. 32. YU GUAN STRUCTURAL DEFINITION
  33. 33. CONCEPT DEVELOPMENT 33 • Systems module • ADCS sensors • Power and Thermal units • Propellant tank • Thermal shielding for cryogenic fuel • Propulsion systems • Escort System • Radiation shield
  34. 34. SYSTEMS MODULE 34
  35. 35. CRYOGENIC PROPELLANT TANK 35 • Internal volume of 625 cubic meters • Aluminum lithium thermal shielding
  36. 36. PROPULSION SYSTEM 36
  37. 37. STRUCTURAL COMPONENTS 37 CYCOM 5320-1 epoxy resin systemfrom Cytek Inc. [12] PAMG-XR1 aerospace grade aluminum honeycomb from Plascore Inc. [13]
  38. 38. RADIATOR DESIGN 38 Hinged radiator used in systems module. [14] Deployable radiator installed in propulsion system. [15]
  39. 39. MASS ESTIMATION 39
  40. 40. STRUCTURAL TESTING 40
  41. 41. YU GUAN COMMUNICATIONS
  42. 42. DESIGN PROCESS  Network Selection  Network ideally suited for needs of the Centurion  Band Selection  Allow for uninterrupted communication without incurring high pointing accuracy requirements  Radiometric Tracking  Enable accurate position and velocity determination of the Centurion  Antenna Selection  Deploy method  Mass  Gain 42
  43. 43. CONCEPT DEVELOPMENT – NETWORK SELECTION  NEN(Near Earth Network) performance is comparable to DSN(Deep Space Network) at L1 and L2 regions(EIRP=85 dBmW)  Fewer missions in L1 and L2 using NEN  Less traffic in wireless communications.  Network of choice: Near Earth Network 43
  44. 44. CONCEPT DEVELOPMENT – FREQUENCY BAND SELECTION Several bands  S-band  Low frequency  Low data rates  Lower pointing requirements  Low atmospheric attenuation  X-band  Frequencies, data rates, pointing requirements, and attenuation in between S and Ka-band  Ka-band  High frequency  High data rates  Higher pointing accuracies  Higher atmospheric attenuation 44 Band Frequency (GHz) S-band 2-3 X-band 7-11 Ka- band 18-30 NEN Frequency Band Characteristics [2] Atmospheric attenuation as a function of frequency [1]
  45. 45. CONCEPT DEVELOPMENT – RADIOMETRIC TRACKING Doppler  Provides velocity estimates  Orbital maneuvers  ΔV calculations Ranging  Provides position estimates  Orbital maneuvers  Orbital transfer points  Docking maneuvers  Get within range of fuel depot and payloads to use proximity sensors 45 Characteristics Value Ranging Accuracy 10 Meters (1 sigma) Doppler Accuracy 1 millimeter per second (1 sigma), 5 second integration time Angle Accuracy 0.1 Degrees Maximum Velocity 2.0 Degrees/second (az and el) Near Earth Network Tracking Characteristics [4]
  46. 46. CONCEPT DEVELOPMENT Parabolic reflector X & S band Low gain as backup Foldable design 46
  47. 47. CRITICAL DESIGN ISSUES  Structural Critical Issues - Thermal Cycling - Material Failure  Communication Critical Issues - Costs of components - Lack of details on ground station 47
  48. 48. SAMIP SHAH ATTITUDE DETERMINATION & CONTROL SYSTEMS
  49. 49. DESIGN PROCESS  Sensor Selection  High attitude sensing requirements for docking and refueling  Actuator Selection  High pointing requirements for docking and refueling  Ability to actively mitigate disturbance torques  Large volume and mass require high performance actuators  Control Systems  Sensor processing and actuator control  Redundancy and flight proven hardware for reliability 49
  50. 50. CONCEPT DEVELOPMENT – SENSOR SELECTION Star Trackers  Absolute attitude sensor  Highest accuracy  Low update rate Selection:  Surrey Rigel-L  2 Units 50Characteristics of common star trackers [22] [23] [24] Surrey Rigel-L Star Tracker Manufacturer/Model Lifetime (yrs) Max Resolution (arc sec) Update Rate (Hz) Tracking Rate (deg/s) Surrey/Rigel-L 7+ X/Y < 3 Z < 25 1-16 6 Terma/HE-5AS N/A RMS pitch, yaw <1 RMS roll <5 4 0.5-2.0 Jena Optronik/ASTRO APS >18 X/Y < 1 Z < 8 10 0.3-3
  51. 51. CONCEPT DEVELOPMENT – SENSOR SELECTION (CONT.) Inertial Measurement Units  Relative attitude sensor  Prone to drift  High update rates Selection:  Honeywell HG9848  2 Units  1 Backup 51 Manufacturer/Mod el Gyro Bias Repeatability/ Stability (deg/h) Gyro ARW (deg/√hr) Gyro Scale Factor (ppm) Accel Bias Repeatability (μg) Scale Factor (ppm) Northrop Grumman/ LN-200S 1/<0.1 <0.07 100 300 300 Honeywell/ HG9848 <0.005 <0.005 <10 <50 150 Kearfott/ KI-4901 0.005/0.003 .003 50 400 500 Characteristics of common IMUs [26] [27] [28]
  52. 52. CONCEPT DEVELOPMENT – SENSOR SELECTION (CONT.) Sun Sensors  Useful for solar array pointing  Relatively inexpensive Selection:  Adcole Course Sun Sensor Pyramid  2 Units 52Characteristics of common sun sensors [5] Adcole Course Sun Sensor Pyramid Manufacturer/Model Field of View Accuracy (deg) # required for full coverage Adcole/ Digital Sun Sensor +/- 64 deg 0.25 5 Adcole/Course Sun Sensor Pyramid 2π steradians 1 2
  53. 53. CONCEPT DEVELOPMENT – SENSOR SELECTION (CONT.) Proximity Sensor  Vital for autonomous docking maneuvers  Enables accurate knowledge of relative position and attitude 53 Distance to Target X (m) Y (m) Z (m) Pitch (deg) Yaw (deg) Roll (deg) Range (m) 2m 0.0003 0.0007 0.0029 0.25 0.09 0.06 0.003 30m 0.0098 0.0039 0.0061 1.04 0.81 0.33 0.012 Demonstrated Accuracy of AOS Proximity Sensors [15] Operation Approach Velocity (m/s) Lateral Alignment (m) Lateral Velocity (m/s) Angular Misalignment (deg) Angular Rate (deg/s) Docking 0.3 0.2 0.05 5 0.25 Berthing 0.01 0.5 < 0.01 < 10 < 0.1 Capture Tolerances for Docking and Berthing [14]
  54. 54. CONCEPT DEVELOPMENT – ACTUATOR SELECTION Attitude Control Thrusters  Provide both large and fine attitude adjustments  16 Units  Clusters of 4  Roll slew time - ~4 min  Pitch/Yaw slew time - ~6 min 54 Manufacturer/Model Thrust (N) Specific Impulse (sec) Propellant Vacco/2 LBF Cold Gas 8.9 - GN2 Moog/DST-11H 22 310 Hydrazine/MON Aerojet/R-1E 111 280 MMH/NTO Aerojet/MR-107V 67 229 Hydrazine Characteristics of common thrusters [31] [32] [33] [34] Configuration of attitude control thrusters
  55. 55. CONCEPT DEVELOPMENT – ACTUATOR SELECTION (CONT.) Control Moment Gyroscopes  Common on large spacecraft  Easily mitigates disturbance torques of roughly 4 x 10-3 N  Reduces thruster firings  Actuate spacecraft in ADCS thruster failure emergency  6 units total  2 units per axis  1 backup per axis  Roll slew time - ~25 min  Pitch/Yaw slew time - ~40 min 55 Manufacturer/Model Angular Momentum (N-m-s) Output Torque (N-m) Weight (kg) Power Requirements (W) Honeywell/M50 25-75 0.075-75 28 95 @ peak torquing L-3/DGCMG 4800/250 4760 258 272 - Airbus/CMG 15-45s 15 45 18.4 25 Characteristics of commonly used control moment gyroscopes [37] [38] [39]
  56. 56. 56
  57. 57. CONCEPT DEVELOPMENT – CONTROL METHODS Onboard Processing  Comprehensive processing is required to control attitude and position  RAD750  Powerful, dependable, flight proven  3 used together for redundancy 57 Characteristics of radiation hardened flight processors [42] [43] [44] Manufacturer/Model Speed (MIPS) Power (W) Flight Proven IBM/Rad6000 35 10 Yes IBM/Rad750 400 10 Yes Proton300k 8000 12 Yes
  58. 58. CONCEPT DEVELOPMENT – CONTROL METHODS (CONT.) 58
  59. 59. SOURCE LINES OF CODE Computer Software Component SLOC Executive 1,000 Communications 2,000 Attitude/Orbit Sensor Processing Sun Sensor 500 IMU 1,000 Star Tracker 2,000 Attitude Determination and Control Kinematic Integration 2,000 Kalman Filter 8,000 Error Determination 1,000 Orbit Propagation 10,000 Attitude Actuator Processing Thruster Control 1,000 CMG Control 1,500 Fault Detection 10,000 Utilities Basic Mathematics 1,000 Transcendental Mathematics 1,500 Matrix Mathematics 2,300 Time Management & Conversion 700 Coordinate Conversion 2,500 Other Functions Momentum Management 3,000 Power Management 2,000 Thermal Control 1,500 Total 54,500
  60. 60. BENTIC SEBASTIAN SPACECRAFT POWER MANAGEMENT
  61. 61. DESIGN PROCESS  Four performance requirements  Supply power for all instruments onboard.  Provide suitable radiation shielding.  Provide suitable temperature for onboard electronics.  Provide backup power when there is no power generation. 61 PROPOSED SOLUTION  ESCORT Bimodal Nuclear Rocket Thermal Propulsion System.  Use radiation hardened instruments in Systems Module, and 3cms of shielding on reactors.  Hinge radiators used to control temperature of Systems Module.  Solar panels during emergency mode
  62. 62. CONCEPT DEVELOPMENT- POWER GENERATION AND DISTRIBUTION 62  Main source of power is BNTRS  Three Nuclear Reactors, producing 25 kW each.  Will be run at 2/3rds of maximum output, 50kW
  63. 63. CONCEPT DEVELOPMENT- POWER GENERATION AND DISTRIBUTION  Peak voltage = 50V for Surrey/Rigel-L star tracker  Peak power input = 113 W for Honeywell/M50 moment gyroscope 63 System component Voltage Required(V) Power Required(W) Surrey/Rigel-L star tracker 16-50V 0.5-6.5W Honeywell/HG9848 IMU 5V 10W Adcole/Coarse Sun Sensor Pyramid 0V 0W Aerojet/R-1E attitude control thruster 28V 36W Honeywell/M50 moment gyroscope 28V 11-113 W IBM/Rad750 2.5-3.3V 5W
  64. 64. POWER GENERATION AND DISTRIBUTION 64
  65. 65. CONCEPT DEVELOPMENT – POWER STORAGE  Requirements for batteries  Light  Rechargeable  High energy density  Long discharging time 65
  66. 66. CONCEPT DEVELOPMENT – POWER STORAGE 66 Technology Specific Density(Wh/kg) Energy Density(Wh/l) Operating temp. Range(C) Design life(years) Cycle life Ag-Zn 100 191 -20 to 25 2 100 Ni-Cd 34 53 -10 to 25 3 25,000-40,000 Super Ni-Cd 28-33 70 -10 to 30 5 58000 IPV Ni-H2 8-24 10 -10 to 30 6.5 At least 60000 CPV Ni-H2 30-35 20-40 -5 to 10 10-14 50,000 SPV Ni-H2 53-54 70-78 -10 to 30 10 At most 30,000 Lithium Ion 90 250 -20 to 30 1 At least 500  To store at least 1.4 kW of power:  Weight of IPV Nickel-Hydrogen batteries: 83 kg  Weight of Lithium batteries(Quallion QL075KA):20 kg
  67. 67. CONCEPT DEVELOPMENT – POWER STORAGE 67  Design choice:  Eight Quallion QL075KA batteries, 72Ah, 3.6V  Eight additional batteries for redundancy.  Total weight of 16 Quallion batteries: 20kg  Advantages of battery.  Quallion QL075KA has a cycle life of 100,000 cycles.  No need to replace batteries.
  68. 68. CONCEPT DEVELOPMENT – RADIATION SHIELDING  Material for Radiation Shielding:  Aluminum  Material Thickness  3cms  Shielding placed in front of Nuclear Reactors 68
  69. 69. CONCEPT DEVELOPMENT – EMERGENCY MODE  Two solar panels will generate additional energy during emergency conditions  Minimum energy required for operations:  1.4kW  Material for Solar Panels:  Amorphous Silicon  Solar panels will have total surface area of 7.7 meters square  Accounting for efficiency of 15%, solar panels provide 23.3 W of power, as emergency power for computers 69
  70. 70. BENTIC SEBASTIAN SPACECRAFT THERMAL SYSTEMS
  71. 71. DESIGN PROCESS  Control temperature within  Nuclear Reactors  Fuel Tank  Systems Module  Solar Panels and Radiators 71 PROPOSED SOLUTION  Two radiators for Nuclear Reactors  Two radiators for Systems Module  Cryogenic fuel tank with MLI and SOFI
  72. 72. CONCEPT DEVELOPMENT – THERMAL CONTROL OF NUCLEAR REACTORS 72 Brayton Cycle of ESCORT System [29]  Peak temperature of HeXe working fluid =929K
  73. 73. CONCEPT DEVELOPMENT – THERMAL CONTROL OF NUCLEAR REACTORS  Deployable radiators designed by Lockheed Martin.  Total area of 65 meters squared.  Will use ammonia coolant loops to control temperatures on radiator.  Radiators will use beta gimbals to keep them perpendicular to the Sun 73
  74. 74. CONCEPT DEVELOPMENT – THERMAL CONTROL OF FUEL TANK 74 Detailed wireframe of the OTV  Cryogenic fuel tank made of Aluminum-Lithium Alloy  60-90 layers of MLI  SOFI with thickness of 30.48 cms
  75. 75. CONCEPT DEVELOPMENT – THERMAL CONTROL OF SYSTEMS MODULE AND SOLAR PANELS 75  Two hinged Radiators of size 2m by 0.5m will be used.  Temperature on solar panels without thermal control can reach upto 846K  Ammonia coolant loop will be used to control temperatures closely. Hinge Radiator by Swales Aerospace[30] Position of radiator and solar panels on inner truss
  76. 76. KEVIN LOHAN LAUNCHING AND DOCKING
  77. 77. DESIGN PROCESS  Launch vehicle  OTV launch vehicle must be reliable  Single launch  Universal docking mechanism  Carry a variety of payloads  Refueling process development 77
  78. 78. CONCEPT DEVELOPMENT : LAUNCH VEHICLE 78 Launch Vehicle Payload Capacity (kg) Launch Cost (millions of $) Falcon XX 140,000 300 Falcon X Heavy 125,000 280 SLS 70,000 500 Falcon 9 Heavy 53,000 85 Delta IV Heavy 22,560 300  US developers  OTV Launch Vehicle  Delta IV Heavy  Custom fairing  Payload and Fuel Launch Vehicle  Falcon 9 Heavy Heavy Class Launch Vehicles [71],[72],[73],[74]
  79. 79. CONCEPT DEVELOPMENT : DOCKING SYSTEM 79  International docking system standards (IDSS)  Regulates where connections are placed  Compatibility for future systems  NASA docking system  Limited active cycles  IDSS compatible NASA Docking System [75] Docking System IDSS Compatibility Probe and Drogue NO APAS NO NASA Docking System YES IDSS compatibility [72] [73]
  80. 80. CONCEPT DEVELOPMENT : REFUELING 80  Robotic Refueling Mission (RRM)  Modified Dextre Arm  Successful test in 2013 Robotic Refueling Mission Arm [78] Refueling System Flow rate (L/min) Time to Refuel (hours) Aerial Refueling 1112 10.5 Gas Pump 37.9 307.8 RRM 1 13725.5 Fluid Flow Rates [60],[61],[62]  Primary Critical Design Issue  Fuel Flow Rate
  81. 81. JAY MULAKALA RISK ANALYSIS
  82. 82. TECHNOLOGY RISK ANALYSIS 82 TECHNOLOGY CONSEQUENCE POF SHORT CODE Star Trackers 3 3 [1] IMU's 3 3 [2] Magnetometer 1 3 [3] Sun Sensor 1 4 [4] Reaction Wheels 2 3 [5] Magnetic Torque Rods 1 1 [6] Control Moment Gyroscopes 4 2 [7] Thrusters 3 4 [8] Nickel-Hydrogen 4 2 [9] Lithium-Hydride 3 2 [10] Solar Panels 3 1 [11] Chemical Thrusters 3 2 [12] Nuclear Thermal Thrusters 5 3 [13] Delta IV 5 2 [14] NASA Docking System 4 3 [15]
  83. 83. OPERATIONAL RISK ANALYSIS 83 OPERATIONS CONSEQUENCE PO SHORT CODE Political setbacks due to technologies 4 4 [1] Use of nuclear power in space 4 4 [2] Issues with nuclear power across nations 4 3 [3] Improper disposal of nuclear fuel 5 1 [4] Accidental failure of nuclear engines 5 2 [5] Lack of funding to complete production 5 1 [6] Improper decommissioning of Centurion 4 2 [7] Launch vehicle failure 2 2 [8] Inclement weather delaying launch 1 5 [9] Launch failure 3 2 [10] Components failing before expected date 1 3 [11] Sabotage 4 2 [12] Human error/negligence 3 3 [13]
  84. 84. 84 RISK MITIGATION Technology Risk Analysis Operational Risk Analysis Main Factors: • Nuclear Thermal Propulsion System • Nickel Hydrogen Batteries Main Factors: • Political setbacks • Nuclear material
  85. 85. JAY MULAKALA MISSION COSTS
  86. 86. COST ESTIMATION 86 FixedCosts • Developmen t • Production • Assembly • $2.518 billion ClientCosts • Fuel Transport • Fuel Costs • $94 million
  87. 87. COST ESTIMATION 87
  88. 88. COST ESTIMATION 88 Centurion’s Liquid Hydrogen Fuel • 10 Missions • 10 Falcon 9 Heavys • $850 million Conventional Bipropellant Fuel • 10 Missions • 50 Falcon 9 Heavys • $4.25 billion About 5 times more over the course of 10 missions
  89. 89. COST ESTIMATION 89 Our Solution • Delta IV • Nuclear bi- modal propulsion system • Modified NASA Docking System Current Solution • Falcon 9 Heavy • Bi-Propellant • 3x Fuel • 2x Cost • NASA Docking System – 2 Cycles
  90. 90. THE CENTURION 90
  91. 91. REFERENCES
  92. 92. REFERENCES
  93. 93. REFERENCES
  94. 94. COST ESTIMATION – FIXED COSTS 94

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